Operation of turbine engines is well known in order to provide propulsion for aircraft.
Referring to FIG. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.
The gas turbine engine 10 operates in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts.
The present invention in particular relates to mounting arrangements for turbine blades in order to provide more efficient cooling of the disc.
FIG. 2 is a schematic perspective view of a prior nozzle guide vane and turbine vane cooling arrangement. The present invention particularly relates to mounting arrangements for turbine blades 118. As can be seen, high pressure coolant air 100 passes through apertures 101 and pre-swirler nozzles 102 into a plenum chamber 103 where it is distributed to root galleries 104 at the base of each blade 118. It will also be understood that coolant air 107, 108, 109 is presented through apertures in guide nozzles 110 and the blades 118 in order to provide surface cooling for these components.
It will be understood that high engine thermal efficiency is dependent upon high turbine entry temperatures which in turn is limited by the material properties of the turbine blades and nozzle guide vane materials. It is for these reasons that cooling is provided within a turbine engine. Furthermore, it will be appreciated that there is heat conduction and convection from the turbine blades 118 to the turbine disc 111 upon which the blades 118 are secured. Typically, in order to achieve cooling, multi-pass coolant systems are provided in which coolant air flows are controlled and regulated through passageways, apertures and nozzles in order to maximise the cooling effect. For example, the pre swirler nozzles 102 act to reduce temperature and pressure of the cooling air flow 100 as it is presented to the rotor disc assembly for cooling purposes.
A particular problem relates to the space, commonly referred to as the wheel space, in front of and behind each rotor disc assembly upon which turbine blades are secured. In short, the air within these spaces has tended to be warmed and is of lower cooling quality due to passing through a rotor stator cavity where it has been exposed to windage and drag from the static components adjacent to the rotor disc. It will be understood for cooling efficiency, it is desirable for there to be a positive flow of relatively cool air. Unfortunately, in some previous arrangements for cooling the blade platform 106 and disc air has been taken from these fore and aft wheel spaces with the result that inefficient cooling has occurred. It will be understood in engines where there is no cooling passage, ingress of air from the relatively hot spaces in front of the blade pocket 105 and its substantial retention and re-circulation within that pocket 105 contributes significantly to high disc 111 temperatures which in turn may limit acceptable gas path temperatures attainable by an engine or reduce effective operational component lifes in the engine.
In order to avoid this particular problem, previous solutions have included adding mechanical features to the front and rear of the disc 111 in order to effectively seal the blade pocket 105 from ingress of the relatively hot wheel space air whilst providing a separate supply of cooling air to cool the blade platform and disc 111 itself. Nevertheless, it will still be appreciated that problems can still arise with high temperatures in the blade pocket 105 leading to higher operating temperatures or restrictions on engine efficiency or limited component life. Furthermore, such mechanical features add to design and assembly complexities as well as weight.
The air flow blowing up the front face of a high pressure turbine disc formed by the front face of that disc and the blade mounting firtree is normally the efflux from a rotary seal which is typically of a labyrinth type. Having flowed through the seal, this airflow impinging upon the front face has been heated and has a tangential velocity in the direction of the rotation of the disc which is less than that of the disc. The result is that there is a raised environmental disc temperature due to the air picking up heat at the seal and the energy input required to bring this air up to disc velocity at the boundary layer with the front face of the high pressure turbine disc.